• Proceeding of EDISON Challenge
• /
• 2012.04a
• /
• pp.77-80
• /
• 2012

초음속 흡입구는 설계점에서 안정적으로 작동하지만 설계점 밖에서는 엔진성능이 급격히 감소하거나 층 격파 불안정 문제가 발생할 수 있다. 초음속 흡입구의 일반적인 특성을 파악하기 위해 2단 꺾임각을 갖는 외부 압축식 2차원 흡입구를 설계하고 EDISON_열유체 시스템을 이용하여 최종적으로 설계 마하수 2.5에서 작동하는 형상을 얻었다. 그러나 설계 마하수 이하의 영역에서는 충격파-경계층, 충격파간 상호작용으로 인해 유동에서 박리가 발생하고 최종적으로 흡입구 목을 질식시켜 아임계 상태로 천이된다. 이를 해결하기 위해 유동 제어 방법 중 하나인 bleeding을 이용하여 경계층을 제거하거나 유동의 박리를 방지하여 충격파를 cowl lip 전방에 안정하게 고정시킬 수 있었으며, 결과적으로 목적하였던 마하수 2.0에서 2.5에 이르는 작동 영역에서 강건하게 운용될 수 있는 초음속 흡입구를 설계하였다.

• Proceeding of EDISON Challenge
• /
• 2015.03a
• /
• pp.585-589
• /
• 2015

본 연구에서는 이중압축램프의 초음속 유동에서 발생하는 충격파 경계층 상호작용을 EDISON_CFD로 해석하기로 한다. 이중압축램프에선 역압력 구배로 인하여 경계층이 박리가 일어나게 되고 박리된 경계층이 다시 이중압축램프에 부착되어 생겨난 박리영역을 관찰할 수 있다. 박리영역의 앞뒤로 유동의 방향이 바뀌게 되면서 압축 팬(compression fan)과 재부착 팬(reattachment fan)이 충격파를 발생시키고 이중압축램프전방의 충격파와 만나서 복잡한 유동 구조를 가지게 됨을 확인하였다. 이와 같은 층류에서 난류, 박리와 재부착의 영역에서의 해석하기 위해선 해석자의 난류모델이 중요하다. $15^{\circ}-30^{\circ}$, $15^{\circ}-45^{\circ}$의 두 종류의 이중압축램프를 $k-{\omega}$ SST 난류 모델과 ${\gamma}-Re_{\theta}$ 천이 모델로 계산을 EDISON_CFD로 수행하였다. 난류 모델의 차이를 표면마찰계수, 압력계수, 마하수로 비교하여 차이점을 분석하였다.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.50 no.11
• /
• pp.747-754
• /
• 2022

An experimental study was carried out to control a shock-induced boundary layer separation by utilizing the supersonic sweeping jet from the fluidic oscillator. High-speed schlieren, surface flow visualization, wall pressure measurement and precise Pitot tube measurement were applied to observe the influences of the location and the supply pressure of the fluidic oscillator on the characteristics of the oblique-shock-induced boundary layer separation. The characteristics of the separation control by the present supersonic fluidic oscillator was quantitatively analyzed by comparing with a conventional control method utilizing an air-jet vortex generator.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.37 no.9
• /
• pp.899-905
• /
• 2009

Most supersonic-flow test facility has axisymmetric nozzles or two-dimensional symmetric nozzles. Compared to these nozzles, a two-dimensional asymmetric nozzle has advantages of reducing low cost for various Mach number testing and undesirable flow structure such as shock wave reflection because the nozzle part can be directly connected to the test section part in this type of nozzle. The two-dimensional asymmetric nozzle, which was Mach number 2, was designed for supersonic combustion experiment. And it was verified with the numerical analysis and visualization of Mach wave. This study suggested the practical method for design and verification of supersonic two dimensional asymmetric nozzles.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2011.11a
• /
• pp.777-780
• /
• 2011

The supersonic wind tunnel is a common facility to studies the aerodynamic phenomenon around the high speed vehicle or weapon system whose operating speed is greater than sonic speed. In this study, a design procedure and selecting the components of a new supersonic wind tunnel whose nozzle exit is $125mm{\times}100mm$ is considered. An operating test of this wind tunnel is being conducted to compare the result with the design values, mach number, etc.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.40 no.12
• /
• pp.1025-1031
• /
• 2012

The present paper deals with concept design of supersonic inlet based on compressible flow theory and flow control of bleeding in order to guarantee stability of supersonic inlet of ramjet engine in broad range of operating conditions. Shock instability, shock wave-boundary layer interaction and flow separation should be properly controlled to improve performance of the supersonic inlet. Considering shock strength, boundary layer and flow separation, the supersonic inlet is modified from the basic model which is designed under inviscid theory. Consequently, shock is stabilized, and required mass flow rate is obtained. Furthermore, bleeding is applied to the supersonic inlet to maintain performance in off-design conditions. Mass flow condition is adopted for modeling of bleeding effect, and performance of the supersonic inlet is evaluated by changing bleeding locations and numbers.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2007.11a
• /
• pp.285-288
• /
• 2007

Numerical study of the influence of the boundary layer of micro-supersonic jet impinging on a flat plate with a hole was performed, to investigate the role of gas jet to eject melted materials from the cut zone in the laser machining. The detailed shock structures and the information of the mass flow rate through the hole were compared to the results of the previous study, in which the effects of boundary layer inside nozzle was not accounted. It was found that the boundary layer inside the micro- nozzle introduced stronger Mach disc over the machining zone, and thus that the mass flow rate through the hole decreased.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.44 no.6
• /
• pp.477-483
• /
• 2016

Velocity profiles of laminar, transition and turbulent boundary layers were investigated by using Particle Image Velocimetry(PIV) measurements on the flat plate at Mach 2.96. The Schlieren visualization and PIV measurements are also used to confirm whether the oblique shock wave generated from the leading edge affects the flow field over the flat plate. The laminar velocity profile measured from the experiment was well matched with the compressible Blasius solution. The velocity profile of the transition boundary layer was well correlated with the theoretical turbulent velocity profile from near the wall and the transition began from Re = $1.41{\times}106$. For the turbulent boundary layer, considering compressibility effects, the Van Driest-transformed velocity satisfies the incompressible log-law. It is found that the log region is extended farther in the wall-normal direction compared to the log region in incompressible boundary layer.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.48 no.3
• /
• pp.207-215
• /
• 2020

Design of a shape transition nozzle is carried out as a part of building a lab-scale supersonic combustion experimental equipment. In order to connect directly the circular shaped vitiation air heater to the square shaped scramjet combustor, area change is evaluated by using the method of characteristics. Shape transition function is introduced to control the transition rate. Boundary layer correction was made through the three-dimensional computational fluid dynamics with the assessment on the several shape transition functions. The shape transition nozzle is proved minimizing the growth of boundary layer at the center of the rectangular nozzle surfaces that caused by the pressure gradient at the corners of the rectangular nozzle and the following recirculation regions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.11a
• /
• pp.387-392
• /
• 2006

A small supersonic wind tunnel was designed and built to study the flow characteristics of a supersonic impulse turbine cascade. Experiments are performed to find flow characteristics of supersonic turbine with the cascade positions and to find factor of expansion loss. The supersonic cascade with a 2-dimensional supersonic nozzle was tested with the cascade positions. Firstly, the flow was visualized by Z-type Schlieren system. Finally, highly complicated flow patterns including shocks, nozzle-cascade interaction and shock boundary layer interactions, flow characteristics of the supersonic turbine were observed.